In operation of a gas turbine engine, air at atmospheric pressure is initially compressed by a compressor and delivered to a combustion stage. In the combustion stage, heat is added to the air leaving the compressor by adding fuel to the air and burning it. The gas flow resulting from combustion of fuel in the combustion stage then expands through a turbine, delivering up some of its energy to drive the turbine and produce mechanical power.
In order to produce a driving torque, the axial turbine consists of one or more stages, each employing one row of stationary nozzle guide vanes and one row of rotating blades mounted on a turbine disc. The nozzle guide vanes are aerodynamically designed to direct incoming gas from the combustion stage onto the turbine blades and thereby transfer kinetic energy to the blades.
The gases typically entering the turbine have an entry temperature from 1850 degrees to 2200 degrees Fahrenheit. Since the efficiency and work output of the turbine engine are related to the entry temperature of the incoming gases, there is a trend in gas turbine engine technology to increase the gas temperature. A consequence of this is that the materials of which the blades and vanes are made assume ever-increasing importance with a view to resisting the effects of elevated temperature.
Historically, nozzle guide vanes and blades have been made of metals such as high temperature steels and, more recently, nickel alloys, and it has been found necessary to provide internal cooling passages in order to prevent melting. It has been found that ceramic coatings can enhance the heat resistance of nozzle guide vanes and blades. In specialized applications, nozzle guide vanes and blades are being made entirely of ceramic, thus, imparting resistance to even higher gas entry temperatures.
However, if the nozzle guide vanes and/or blades are made of ceramic, which have a different chemical composition, physical property and coefficient of thermal expansion to that of a metal supporting structure, then undesirable stresses, a portion of which are thermal stresses, will be set up between the nozzle guide vanes and/or blades and their supports when the engine is operating. Such undesirable thermal stresses cannot adequately be contained by cooling.
Furthermore, conventionally the interface between a turbine's metallic blades and the rigid turbine stator will accept a predetermined amount of rubbing therebetween. However, ceramic blade tips cannot rub as have conventional metallic blades. The result of ceramic blade rubbing with the rigid turbine stator is catastrophic failure. Thus, an increased blade tip clearance is required to avoid tip rubbing resulting in a significant performance penalty.
The present invention is directed to overcome one or more of the problems as set forth above.